Spacecraft of all types are subject to various undesirable interactions with their ambient space plasma environment. Many geosynchronous and polar orbiting satellites, for example, experience operational anomalies caused by spacecraft charging. Many satellite failures are attributable to the failure of their solar arrays, and at least half of the solar array failures are believed to be charging induced. In general, two sources of spacecraft charging dominate as causes for generating large differences in potential on spacecraft surfaces: (1) high-energy sub-storm electron currents, and (2) photo-emission currents. During geomagnetic sub-storms, all surfaces of a synchronous orbit spacecraft collect excess high-energy electrons (>>10 keV) and consequently charge negatively. However, the sunlit surfaces of the spacecraft continuously emit photoelectrons and consequently approach plasma potential. The opposing charging phenomena can cause a large potential difference to develop between the sunlit and dark surfaces. This potential difference can lead to catastrophic electrostatic discharging.
FIG. 1 illustrates the process with a satellite 10, having two wings of solar arrays (or panels) extending from opposite sides of the satellite's body. One side 2 of each solar array faces the sun, while another side 4 lies in shadow. On the sunny side, plasma ions 6 and plasma electrons 8 impinge upon sunlit surfaces. Some incident electrons reflect off the sheathing or cover glass of the solar array, other electrons backscatter after reaching the solar array surface, while solar photons 14 induce photoelectrons 12 to leave the array surfaces. The positive charging of the array surface repels some plasma ions 6, while some photoelectrons 12 return to the surface, attracted by the positive surface charge. Photoemission typically produces a current density in excess of one nA/cm2. Meanwhile, on the shadowed side 4, plasma ions 6 and plasma electrons 8 impinge upon dark surfaces, with the developing negative charge of the surface repelling some plasma electrons.
The unequal charging of different surfaces of the spacecraft results in inverted electric-field gradients, as illustrated in FIG. 2. The structure (or chassis) 16 of the spacecraft 10 charges negatively, while the sunlit front surfaces of the solar arrays 18 charge positively. The potential difference that develops between the surfaces of the arrays and the chassis can range from hundreds to thousands of volts, and can increase over time. The graph 20 of FIG. 3 shows an increasing potential difference between a sunlit surface and the spacecraft chassis. In the graph, the y-axis represents potential in volts, and the x-axis represents time. Plot 22 represents the charging over time, for example, of a sunlit cover of a solar cell, whereas plot 24 represents the charging over time of the spacecraft chassis. The separation between the plots at any given time indicates the difference in potential between the surfaces. For example, at time 0, the voltages are equal, but after 500 seconds of charging, a voltage difference approximating 1000 volts develops because of the charging phenomena described above, with the solar cell surface charging less negatively than the chassis.
The threat posed by inverted gradients is that electrostatic discharge (ESD) can occur at lower potentials than normal gradient charging (on the scale of hundreds of volts, instead of thousands). FIG. 4A, FIG. 4B, and FIG. 4C show a destructive process that can result from the inverted electric-field gradients that form between a sunlit surface and the spacecraft chassis. In FIG. 4A, a solar cell cover 30 shields the underlying interconnections 32 of the solar array. The solar cell array sits on a conductive chassis or frame 34 (i.e., ground). The interconnections 32 are generally small conductive traces or wires that string together the solar cells of an array. The cover 30 is typically made of a dielectric material, effectively insulated from ground. Differential charging between adjacent surfaces (e.g., between solar array strings) can induce a primary arc of ESD. This primary arc generates a plasma cloud 36.
In FIG. 4B, the plasma cloud 36 provides a conduction path 38 between solar array strings 30 at different potentials. An actively powered solar array string provides the power to sustain high-energy secondary arcing or discharges. These secondary arcs may result in catastrophic permanent array damage by opening or short-circuiting adjacent solar array strings or power conditioning components. FIG. 4C shows, for example, a carbon track 40 amidst the interconnections 32 that shorts neighboring array strings. This problem pervades the satellite industry.